1. Field of the Invention
The present invention relates to the field of multispool, notably twin-spool, turbojets and more generally to that of turbine engines.
2. Description of the Related Art
A twin-spool turbojet comprises a gas turbine engine through which travels a gas flow called the main flow driving a fan delivering an airflow called the bypass flow. When the fan is placed in front of the engine, it discharges the aspirated air that is separated into two concentric flows: one being the main flow and the other being the bypass flow. The air of the main flow is compressed again and then mixed with a fuel in a combustion chamber in order to produce a gaseous flow with high energy which sets in motion the turbines placed downstream. One of the turbines is connected by a shaft to the fan rotor that it drives. The bypass flow in civil aircraft engines supplies most of the thrust of the engine and the diameter of the fan is consequently very large.
The fan rotor comprises an impeller the hub of which is secured to the drive shaft and the rim comprises blade slots oriented substantially axially. The axial direction is that of the engine shaft. The blades are engaged by their root in the individual slots and form the fan rotor. A fan blade comprises a root, an airfoil with an aerodynamic profile and a stilt between the root and the airfoil. In order to form the border surface between the rim of the rotor and the airstream and to ensure the continuity of the main stream, intermediate platforms are placed between the blades. Being different from the upper compression stages and because of the considerable dimensions of the blades, the fan rotor platforms do not form an integral part of the blades but are separate parts. It should be noted that the internal radius of the airstream increases notably between the inlet and the outlet of the fan rotor.
A clearance is arranged between the platforms and the blades in order to allow the latter a limited range of movement during the various operating phases of the engine. This clearance is plugged by a seal of elastomer attached along the lateral edges of the platform and resting against the adjacent blade.
According to the prior art, the seal is of elongate shape with a constant profile from one end to the other. Transversely, it consists of three portions: a portion for attachment to the platform, a flexible portion and a bulbous portion formed to ensure a contact with the surface of the adjacent part. The flexible portion allows the seal to be suited to the distance separating the edge of the platform from the surface of the facing blade.
It is noted that, after a certain period of operation of the turbine engine on which they are mounted, they have zones of wear and of breakage. The result of this is a loss of seal at the fan blade root. A poor seal has an impact on the flow rate and the efficiency of the compressor stages directly downstream of the fan. It also has an impact on the margin of surging.
The seals are therefore parts that need to be replaced regularly throughout the lifetime of the turbine engine to ensure optimum operation of the latter.